EP1621733B1 - Conduit de transition pour une turbine à gaz - Google Patents

Conduit de transition pour une turbine à gaz Download PDF

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Publication number
EP1621733B1
EP1621733B1 EP05015705.6A EP05015705A EP1621733B1 EP 1621733 B1 EP1621733 B1 EP 1621733B1 EP 05015705 A EP05015705 A EP 05015705A EP 1621733 B1 EP1621733 B1 EP 1621733B1
Authority
EP
European Patent Office
Prior art keywords
flow
ribs
supporting ribs
guide
channel
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Fee Related
Application number
EP05015705.6A
Other languages
German (de)
English (en)
Other versions
EP1621733A3 (fr
EP1621733A2 (fr
Inventor
Martin Dr. Hoeger
Franz Malzacher
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by MTU Aero Engines AG filed Critical MTU Aero Engines AG
Publication of EP1621733A2 publication Critical patent/EP1621733A2/fr
Publication of EP1621733A3 publication Critical patent/EP1621733A3/fr
Application granted granted Critical
Publication of EP1621733B1 publication Critical patent/EP1621733B1/fr
Expired - Fee Related legal-status Critical Current
Anticipated expiration legal-status Critical

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/545—Ducts
    • F04D29/547—Ducts having a special shape in order to influence fluid flow
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to a flow structure for a transition duct of a gas turbine, according to the preamble of patent claim 1.
  • Gas turbines such as aircraft engines, typically have multiple compressors, multiple turbines, and a combustor.
  • the multiple compressors are usually a low-pressure compressor and a high-pressure compressor, with the multiple turbines being a high-pressure turbine and a low-pressure turbine.
  • the gas turbine is flowed through in the axial direction, wherein the low-pressure compressor upstream of the high-pressure compressor and the high-pressure turbine is positioned upstream of the low-pressure turbine. From the low pressure compressor, the flow enters the high pressure compressor through a transitional passage between these two compressors. Likewise, such a transitional channel is positioned between the high pressure turbine and the low pressure turbine.
  • support ribs spaced apart from one another in such transitional channels between two compressors or two turbines in the circumferential direction of the transition duct.
  • the support ribs serve to carry out, for example, oil lines and sensors and the absorption of forces, which is why the support ribs are made relatively thick.
  • Support ribs are known from the prior art, although they are designed to flow but not flow diverting.
  • support ribs are already known which have a suction side and a pressure side and therefore also assume the function of a flow deflection.
  • a flow structure for a transition duct of a gas turbine is known in which a plurality of guide ribs are positioned between each two adjacent support ribs, wherein the flow exit edges of the guide ribs run upstream of the flow exit edges of the support ribs. This structure can reduce the risk of flow separation.
  • the present invention is based on the problem to provide a further improved flow structure for a transitional channel of a gas turbine.
  • the flow entry edge of the or each guide rib positioned between two adjacent support ribs extends upstream of the flow entry edges of the support ribs.
  • a transition wall radially inwardly delimiting the channel wall and / or the transition channel radially outwardly delimiting channel wall in the region of the flow outlet edge of the or each, positioned between two adjacent support ribs guide rib is retracted inward.
  • Fig. 2 can be removed, it is already known from the prior art, in the region of the transitional channel 10 a plurality of circumferentially of the transitional channel 10 spaced-apart support ribs 16 to be arranged.
  • the support ribs are relatively thick and relatively long and have a small height ratio.
  • Fig. 2 It is already known from the prior art to form the support ribs 16 as flow-deflecting support ribs and therefore to contour their side surfaces in the sense of a pressure side 17 and a suction side 18.
  • the flow separations in Fig. 2 are visualized by arrows 19 and the secondary flows by arrows 20.
  • Such flow separation and secondary flows affect the efficiency of the gas turbine and are therefore disadvantageous overall.
  • Fig. 3 to 5 show a preferred embodiment of a flow structure according to the invention for a gas turbine in the region of a transitional channel 21 between a medium-pressure compressor 22 and a low-pressure compressor and a high-pressure compressor 23, wherein Fig. 3 the medium-pressure compressor 22 positioned in the flow direction in front of the transition duct 21 is in turn closed by a rotor blade ring 24 in the region of its last stage.
  • the blade ring 24 comprises according to Fig. 4 a plurality of circumferentially spaced apart blades 25, wherein the direction of rotation of the blade ring 24 and the blades 25 is visualized by an arrow 26.
  • spaced apart support ribs 27 and 28 are arranged in the circumferential direction, wherein the support ribs 27 and 28 have a pressure side 29 and a suction side 30 and are therefore designed flow diverting.
  • the flow outlet edges 32 of the guide ribs 31 extend in a region which, starting from a flow inlet edge 34 of the support ribs 27 and 28, lies between 30% and 50% of the chord length of the support ribs 27 and 28.
  • the guide ribs 31 are therefore formed relatively short relative to the support ribs 27 and 28.
  • the guide ribs 31 are positioned between the support ribs 27 and 28 such that flow entry edges 35 of the guide ribs 31 are located slightly upstream of the flow entry edges 34 of the support ribs 27 and 28.
  • the flow inlet edges 34 of the support ribs 27 and 28 extend from flow inlet edges 35 of the guide ribs 31 in a range between 5% and 10% of the chord length of the guide ribs 31st
  • transition channel 21 is limited on the hub side by a radially inner channel wall 36 and the housing side of a radially outer channel wall 37.
  • the channel walls 36 and 37 are now contoured such that they are drawn inward in the region of the flow outlet edge 32 of the guide ribs 31 while narrowing the flow cross-section.
  • FIGS. 3 and 5 Dashed contour shown.
  • Fig. 5 can be taken that the channel walls 36 and 37 adjacent to the suction side 30 are less retracted inwardly as adjacent to the pressure side 29 of two adjacent support ribs 27, 28. This creates a circumferentially non-rotationally symmetrical contour of the channel walls 36 and 37, which in terms of a flow optimization is particularly preferred.
  • the guide ribs 31 being relatively short and relatively thin with respect to the support ribs 27, 28.
  • the flow inlet edges 35 of the guide ribs 31 are arranged slightly upstream of the flow inlet edges 34 of the support ribs 27, 28, but the flow outlet edges 32 of the guide ribs 31 extend significantly upstream of the flow outlet edges 33 of the support ribs 27, 28.
  • the chord length of the support ribs 31 is in about 30% 50% of the chord length of the support ribs 27, 28.
  • channel walls 36 and 37 of the transition channel 21 are retracted inwards. This creates a circumferential non-rotationally symmetrical side wall contour of the transition channel.
  • the flow structure according to the invention is used either between two compressors or two turbines or at the turbine outlet housing downstream of the low-pressure turbine.
  • the use of the flow structure according to the invention in a transitional channel between a medium-pressure compressor and a high-pressure compressor or in a transitional channel between a high-pressure turbine and a low-pressure turbine and downstream of the low-pressure turbine of an aircraft engine is preferred.
  • the use of the flow structure is particularly preferred when a medium-pressure compressor or compressor positioned upstream of the transition duct is closed off at the downstream stage by a blade ring, as seen in the flow direction. As a result, then the length can be shortened.
  • the flow structure according to the invention can be produced relatively inexpensively as a cast component. Flow losses in the transition channel are significantly reduced. It is possible to save the so-called exit guide wheel in the upstream compressor, the exit guide wheel in the upstream low-pressure turbine or the inlet guide wheel into the low-pressure turbine. This results in a more compact and lighter design. The efficiency of a gas turbine can therefore be effectively increased.

Claims (7)

  1. Structure d'écoulement pour une canal de transition (21) d'une turbine à gaz avec des nervures d'appui (27, 28) espacées l'une de l'autre dans le sens périphérique du canal de transition (21), positionnées dans le canal de transition (21), dans laquelle entre respectivement deux nervures d'appui (27, 28) contigües, espacées l'une de l'autre dans le sens périphérique du canal de transition (21), au moins une pale de guidage ou nervure de guidage (31) est positionnée, et dans laquelle l'arête de sortie d'écoulement (32) de la ou chaque nervure de guidage (31) s'étend en amont des arêtes de sortie d'écoulement (33) des nervures d'appui (27, 28),
    caractérisée en ce que
    l'arête d'entrée d'écoulement (35) de la ou chaque nervure de guidage (31) s'étend en amont des arêtes d'entrée d'écoulement (34) des nervures d'appui (27, 28).
  2. Structure d'écoulement selon la revendication 1,
    caractérisée en ce que
    l'arête de sortie d'écoulement (32) de la ou chaque nervure de guidage (31) s'étend à partir des arêtes d'entrée d'écoulement (34) des nervures d'appui (27, 28) dans une zone entre 30 jusqu'a 50% de la longueur de corde des nervures d'appui (27, 28).
  3. Structure d'écoulement selon la revendication 1 ou 2,
    caractérisée en ce que
    les arêtes d'entrée d'écoulement (34) des nervures d'appui (27, 28) s'étendent à partir d'arêtes d'entrée d'écoulement (35) des nervures de guidage (31) dans une zone entre 5 jusqu'a 10 % de la longueur de corde des nervures de guidage (31).
  4. Structure d'écoulement selon l'une quelconque des revendications 1 à 3,
    caractérisée en ce que
    les nervures de guidage (31) sont réalisées par rapport aux nervures d'appui (27, 28) relativement courtes et mincies.
  5. Structure d'écoulement selon l'une quelconque des revendications 1 à 4,
    caractérisée en ce que
    les nervures d'appui (27, 28) sont réalisées de manière à renvoyer l'écoulement avec un côté d'aspiration (30) et un côté de refoulement (29).
  6. Structure d'écoulement selon l'une quelconque des revendications 1 à 5,
    caractérisée en ce que
    une paroi de canal (36) délimitant le canal de transition (21) radialement vers l'intérieur et/ou une paroi de canal (37) délimitant radialement vers l'extérieur le canal de transition dans la zone de l'arête de sortie d'écoulement (32) de la ou chaque nervure de guidage (31) positionnée entre deux nervures d'appui (27,28) contigües, est intégrée vers l'intérieur.
  7. Structure d'écoulement selon la revendication 6,
    caractérisée en ce que
    la paroi de canal intérieure (36) et/ou la paroi de canal extérieure (37) est intégrée plus fortement entre deux nervures d'appui contigües (27, 28) dans la zone d'un côté de refoulement (29) d'une nervure d'appui (28) que dans la zone d'un côté d'aspiration (30) de la nervure d'appui contigüe (27).
EP05015705.6A 2004-07-28 2005-07-20 Conduit de transition pour une turbine à gaz Expired - Fee Related EP1621733B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102004036594A DE102004036594A1 (de) 2004-07-28 2004-07-28 Strömungsstruktur für eine Gasturbine

Publications (3)

Publication Number Publication Date
EP1621733A2 EP1621733A2 (fr) 2006-02-01
EP1621733A3 EP1621733A3 (fr) 2011-12-21
EP1621733B1 true EP1621733B1 (fr) 2017-03-01

Family

ID=35079143

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05015705.6A Expired - Fee Related EP1621733B1 (fr) 2004-07-28 2005-07-20 Conduit de transition pour une turbine à gaz

Country Status (3)

Country Link
US (1) US7553129B2 (fr)
EP (1) EP1621733B1 (fr)
DE (1) DE102004036594A1 (fr)

Families Citing this family (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB0624294D0 (en) * 2006-12-05 2007-01-10 Rolls Royce Plc A transition duct for a gas turbine engine
US8224432B2 (en) * 2007-05-08 2012-07-17 C.R.Bard, Inc. Rapid 3D mapping using multielectrode position data
US8182204B2 (en) * 2009-04-24 2012-05-22 Pratt & Whitney Canada Corp. Deflector for a gas turbine strut and vane assembly
US10221707B2 (en) 2013-03-07 2019-03-05 Pratt & Whitney Canada Corp. Integrated strut-vane
US9879540B2 (en) * 2013-03-12 2018-01-30 Pratt & Whitney Canada Corp. Compressor stator with contoured endwall
US9835038B2 (en) 2013-08-07 2017-12-05 Pratt & Whitney Canada Corp. Integrated strut and vane arrangements
US9556746B2 (en) 2013-10-08 2017-01-31 Pratt & Whitney Canada Corp. Integrated strut and turbine vane nozzle arrangement
FR3032495B1 (fr) * 2015-02-09 2017-01-13 Snecma Ensemble de redressement a performances aerodynamiques optimisees
FR3034820B1 (fr) * 2015-04-13 2019-07-12 Safran Aircraft Engines Piece de turbomachine a surface non-axisymetrique
GB201512838D0 (en) 2015-07-21 2015-09-02 Rolls Royce Plc A turbine stator vane assembly for a turbomachine
US9909434B2 (en) 2015-07-24 2018-03-06 Pratt & Whitney Canada Corp. Integrated strut-vane nozzle (ISV) with uneven vane axial chords
US10443451B2 (en) 2016-07-18 2019-10-15 Pratt & Whitney Canada Corp. Shroud housing supported by vane segments
FR3059735B1 (fr) * 2016-12-05 2020-09-25 Safran Aircraft Engines Piece de turbomachine a surface non-axisymetrique
GB201703422D0 (en) 2017-03-03 2017-04-19 Rolls Royce Plc Gas turbine engine vanes
GB201703423D0 (en) 2017-03-03 2017-04-19 Rolls Royce Plc Gas turbine engine vanes
DE102017212311A1 (de) 2017-07-19 2019-01-24 MTU Aero Engines AG Umströmungsanordung zum Anordnen im Heißgaskanal einer Strömungsmaschine
GB2568109B (en) * 2017-11-07 2021-06-09 Gkn Aerospace Sweden Ab Splitter vane
US11396888B1 (en) * 2017-11-09 2022-07-26 Williams International Co., L.L.C. System and method for guiding compressible gas flowing through a duct
DE102017222193A1 (de) 2017-12-07 2019-06-13 MTU Aero Engines AG Turbomaschinenströmungskanal
DE102017222817A1 (de) 2017-12-14 2019-06-19 MTU Aero Engines AG Turbinenmodul für eine strömungsmaschine
US11560797B2 (en) * 2018-03-30 2023-01-24 Siemens Energy Global GmbH & Co. KG Endwall contouring for a conical endwall
US11859515B2 (en) * 2022-03-04 2024-01-02 General Electric Company Gas turbine engines with improved guide vane configurations

Family Cites Families (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2847822A (en) * 1954-11-01 1958-08-19 United Aircraft Corp Thrust modifying device
US2990106A (en) * 1956-10-12 1961-06-27 English Electric Co Ltd Axial flow multi-stage compressors
US3704075A (en) * 1970-12-14 1972-11-28 Caterpillar Tractor Co Combined turbine nozzle and bearing frame
DE2741063C2 (de) * 1977-09-13 1986-02-20 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Gasturbinentriebwerk
GB2115881A (en) * 1982-02-26 1983-09-14 Rolls Royce Gas turbine engine stator vane assembly
US4989406A (en) * 1988-12-29 1991-02-05 General Electric Company Turbine engine assembly with aft mounted outlet guide vanes
US5316441A (en) * 1993-02-03 1994-05-31 Dresser-Rand Company Multi-row rib diffuser
US6375419B1 (en) * 1995-06-02 2002-04-23 United Technologies Corporation Flow directing element for a turbine engine
GB9805030D0 (en) * 1998-03-11 1998-05-06 Rolls Royce Plc A stator vane assembly for a turbomachine
GB9823840D0 (en) * 1998-10-30 1998-12-23 Rolls Royce Plc Bladed ducting for turbomachinery
SE0004001D0 (sv) * 2000-11-02 2000-11-01 Atlas Copco Tools Ab Axial flow compressor
DE10210866C5 (de) * 2002-03-12 2008-04-10 Mtu Aero Engines Gmbh Leitschaufelbefestigung in einem Strömungskanal einer Fluggasturbine
US7094027B2 (en) * 2002-11-27 2006-08-22 General Electric Company Row of long and short chord length and high and low temperature capability turbine airfoils
US6905303B2 (en) * 2003-06-30 2005-06-14 General Electric Company Methods and apparatus for assembling gas turbine engines

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None *

Also Published As

Publication number Publication date
EP1621733A3 (fr) 2011-12-21
US7553129B2 (en) 2009-06-30
US20060024158A1 (en) 2006-02-02
DE102004036594A1 (de) 2006-03-23
EP1621733A2 (fr) 2006-02-01

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